Next-generation spacecraft will use star trackers and Micro Electro-Mechanical System (MEMS) gyros for attitude determination and control. MEMS gyros are solid-state devices that are used widely in consumer products and are now being adapted for space applications. They are manufactured in large quantities using processes from the semi-conductor industry, are very small, low power, light weight, and inherently low cost. MEMS gyros may be packaged with a star tracker to create a hybrid sensor that provides both attitude and rate data. Using hybrid sensors eliminates the need for a high performance Inertial Measurement Unit (IMU) that can cost up to $3 million and weigh up to 45 lbs. The disadvantage of MEMS gyros is that they have much lower accuracy than gyros used today in precision spacecraft IMUs (e.g., 100 times or greater performance reduction). However, because star trackers can provide continuous three-axis attitude sensing, MEMS gyro data is only needed for brief periods when star tracker data is unavailable. Such data outages may occur when a star tracker is subjected to Sun interference, or is obstructed by the Earth, or during failure recovery and safe mode events.
A challenge with a MEMS based attitude determination (AD) system is getting the required pointing accuracy when the spacecraft must operate in the absence of star tracker data. As is the case with other inertial measurement devices, MEMS may have large temperature sensitivities, such that their performance varies significantly as temperature conditions change. One solution, typically employed for precision IMUs, is to actively control the MEMS gyros temperatures. However, when MEMS gyros are packaged with a star tracker, maintaining temperature control requires biasing the MEMS gyro temperatures well above the star tracker baseplate temperature. Unavoidable heat leaks with this approach increase the star tracker operating temperature and reduce star-sensing accuracy. Alternatively, the MEMS gyros temperature sensitivities can be calibrated on the ground and applied in orbit over the mission life. The drawback of this approach is that these temperature sensitivities may change significantly over the in-orbit life (typically 15 years) due to aging and radiation effects.